• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.7-10
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• 2009

The development and evaluation of solid propellant were performed for the turbopump pyro starter, which start up the liquid propellant rocket engine for the Space Launch Vehicle (SLV). Requirements for the turbopump pyro starter propellant include the production of low flame temperature, low burning rate and nontoxic gas to protect the mechanical corrosion or air pollution. This study describes the development of the solid propellant composition which is based on PCP binder. DHG (Dihydroxy glyoxime), which has advantages of oxygen balance and ignition, was used as coolant. The mechanical properties and burning rate of the propellants were measured. Finally, static fired test was performed to prove the possibility of development.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.94-97
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• 2008

As the partial admission turbine has a intrinsically unsteady and three dimensional flow region, numerical calculation time of these study has been too long time. The numerical analysis for gas temperature of turbine blade surface has been performed to investigate development of temperature with various pyro start pressure. Computations have been carried out several turbine rotational speeds in the range from 0 to 16000 rpm and inlet conditions with 1423K, 7.2MPa. As a result, the more rotational speed and pyro starter pressure of turbine increased, the more turbine blade's temperature decreased. It is also found that flow field of turbine blade inlet area at pyro starter pressure of 5.75MPa and rotational speed of 12100 rpm formed surface temperature uniformly.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.167-170
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• 2006

For a short time a pyre-starter should turn the blades of a turbine to the adequate rotational speed by a single operation. Through this process the pressures of the components of a propellant rise rapidly up to the operating point, and the components enter into a gas-generator. Combustion in the gas-generator occurs to keep the turbopumps working. In this research characteristic parameters of a pyre-starter which correspond to the required performance of the turbopump before the gas-generator starts to work were selected

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.35-41
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• 2013

Torque characteristics of a turbopump turbine was analyzed using the turbine performance test result. Specific torque of the subject turbine could be expressed as a linear function of corrected rotor speed at a fixed pressure ratio and it has been confirmed by the test result. It also found that corrected rotor speed-specific torque characteristics does not change anymore if the turbine pressure ratio is set bigger than a certain value. Using the turbine torque characteristics and pyro-starter performance test results, rotational speed development behavior of the turbopump was predicted. Prediction revealed that the lap time reaching 50% of turbopump design speed is less than 0.7 second. Effect of the thermal loss between pyro-starter and turbopump was negligible.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.4-10
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• 2012

Torque characteristics of a 75-tonf turbopump turbine was analyzed using the turbine performance test result. Specific torque of the subject turbine could be expressed as a linear function of corrected rotor speed at a fixed pressure ratio and it has been confirmed by the test result. It also found that corrected rotor speed-specific torque characteristics does not change anymore if the turbine pressure ratio is set bigger than a certain value. Using the turbine torque characteristics and pyro-starter performance test results, rotational speed development behavior of the turbopump was predicted. Prediction revealed that the lap time reaching 50% of turbopump design speed is less than 0.7 second. Effect of the thermal loss between pyro-starter and turbopump was negligible.

• Aerospace Engineering and Technology
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• v.7 no.2
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• pp.170-175
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• 2008

In this study, it was investigated that the characteristics of startup and compatibility using several type hot and cold gases. The characteristics of starting LRE by pyro starter was compared with that by a Helium spinner. The compatibility of pyro gas, a gaseous Helium, Hydrogen+Nitrogen mixture gas, and air was investigated by a simple 1D turbine analysis considered the properties of each gas and turbine efficiency. Most of them were compatible to start up the LRE however air was properly used only for low power mode of turbine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.623-626
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• 2009

Hanwha Corporation Daejeon Plant have developed apogee Kick Motor of KSLV-I that is the first among nation space launch vehicle for five years from 2003. Now, we are joining in KSLV-II(Korea Space Launch Vehicle-II) project and developing Pyro starter which is turbo pump for the first start-up of liquid propulsion supply.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.216-220
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• 2006

In this study, it was investigated that the characteristics of startup and compatibility using several type hot and cold gases. The characteristics of starting LRE by pyro starter was compared with that by a He spinner. The compatibility of pyre gas, a gaseous He, H2+N2 mixture gas, and air was investigated by a simple 1D turbine analysis considered the properties of each gases and turbine efficiency. Most of them were compatible to start up the LRE but air was properly used only when the turbine was low power mode.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.2
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• pp.74-80
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• 2008

The feasibility study into the development of turbine spinners, which start up the turbo-pump, has been carried out and the design requirements and parameters ranges have been presented. Turbine spinners use the solid propellant as such composite propellant based AN compound with high energy plasticizers, coolants, and phase stabilizer which relieves a sensible volume change due to the phase transformation of AN near room temperature. Propellants which have a homing rate of $0.2{\sim}0.3\;mm/s$ and pressure exponent ranged from 0.3 to 0.6, showed stable burn-out in the standard motor tests. Both the magnitude of ignition energy and its thermal transfer mechanism have been proved to have a tangible effect on the ignition of the pyre starter, and the results of this study showed that a flame temperature of 1400K would be quite adequate to get a stable ignition for the AN composite propellant.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.4
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• pp.48-55
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• 2008

In this paper, numerical analysis of the surface gas temperature on turbine blades has been performed to investigate the temperature profiles characteristics of a partial admission supersonic turbine driven by high temperature and pressure gas of pyro-starter with two different types of turbine blade edge shape. In order to examine the surface gas temperature on turbine blades at initial starting, computations tlave been carried out at several turbine rotational speeds in the range of $0{\sim}10,000$ rpm for each type of turbine edge shape. "Sharp" edge and "Round" edge types were taken as the turbine edge shape factor. As turbine rotational speed increased, the average temperature of turbine blades was further decreased. It was also found that the surface temperature of turbine blades with a sharp edge was lower than round-type edge turbine blades.